US7484937B2 - Compressor blade with reduced aerodynamic blade excitation - Google Patents

Compressor blade with reduced aerodynamic blade excitation Download PDF

Info

Publication number
US7484937B2
US7484937B2 US10/870,437 US87043704A US7484937B2 US 7484937 B2 US7484937 B2 US 7484937B2 US 87043704 A US87043704 A US 87043704A US 7484937 B2 US7484937 B2 US 7484937B2
Authority
US
United States
Prior art keywords
blade
leading edge
boundary layer
layer flow
suction side
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Fee Related
Application number
US10/870,437
Other versions
US20050271513A1 (en
Inventor
Erik Johann
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Family has litigation
First worldwide family litigation filed litigation Critical https://patents.darts-ip.com/?family=34928806&utm_source=google_patent&utm_medium=platform_link&utm_campaign=public_patent_search&patent=US7484937(B2) "Global patent litigation dataset” by Darts-ip is licensed under a Creative Commons Attribution 4.0 International License.
Application filed by Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: JOHANN, ERIK
Publication of US20050271513A1 publication Critical patent/US20050271513A1/en
Application granted granted Critical
Publication of US7484937B2 publication Critical patent/US7484937B2/en
Anticipated expiration legal-status Critical
Expired - Fee Related legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/16—Form or construction for counteracting blade vibration
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/04—Antivibration arrangements
    • F01D25/06—Antivibration arrangements for preventing blade vibration
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D21/00—Pump involving supersonic speed of pumped fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/14—Two-dimensional elliptical
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/16—Two-dimensional parabolic
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/60—Structure; Surface texture
    • F05D2250/61—Structure; Surface texture corrugated
    • F05D2250/611—Structure; Surface texture corrugated undulated
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to a compressor blade with reduced aerodynamic blade excitation, in particular for the fan of aircraft engines.
  • Compressor blades with a large chord length used on certain types of aircraft-engine compressors are loaded to such an extent under certain operating conditions that damage to or life reduction of the compressor blade and the compressor disk may occur.
  • the local fixation of the transition area away from the compression shock area suppresses, or at least limits, coupling of the interrelated oscillation in the transition and compression shock area and, ultimately, avoids oscillation of the compression shocks or limits them to such an extent that their vibration-augmenting reaction on the natural vibrations of the compressor blade occurring under certain flight conditions, which eventually may lead to damage of the compressor blades and the compressor disk, is prevented.
  • a broad aspect of this invention is to provide a design of compressor blades such that, independently of the already suggested fixation of the flow transition point by applying a surface roughness to the suction side, oscillation of the compression shocks (suction side) and their vibration-augmenting coupling to the natural vibrations of the blade is avoided or reduced, in order to avoid damage to the compressor disks and blades occurring under certain operating conditions.
  • the essence of this invention is the particular design of the blade leading edge such that the leading edge shock wave effective there will not detach from, but rather attach to the leading edge to effect the transition from the laminar to the turbulent boundary layer flow at a short distance from the leading edge without accelerating it along the continuous curvature of the suction side starting at the leading edge, thus keeping the boundary layer flow constant without re-lamination. Accordingly, the transition point, whose periodic movement is largely suppressed and which is sufficiently remote of the compression shock on the suction side, can neither communicate with nor have a vibration-augmenting effect on the latter.
  • reaction of the vibrations of the compression shock onto the critical natural vibrations occurring under certain flight conditions is suppressed or at least limited to such an extent that the natural vibrations are not augmented, or are not augmented to such an extent that overstressing in excess of the strength limits, with consequential material damage to the compressor blades and/or the compressor disk, occurs.
  • the shape of the blade leading edge can, for example, be a very slender ellipse or parabola—with an ellipse ratio of the semi-minor axis to the semi-major axis smaller than or equal to 1:4—with the curvature extending continuously on the suction side.
  • the leading edge shock wave attaches to the leading edge, the initially laminar boundary flow quickly transforms, i.e. at a short distance from the blade leading edge, into a turbulent boundary flow without being accelerated and re-laminated, and the transition point is thus not moved periodically.
  • the leading edge can, in the upper, natural-vibration critical area, also have a serrated structure, pocket-type depressions or a recessed area acting as sweep-back. Additionally, other designs of the blade leading edge are possible to ensure attachment of the leading edge shock wave to the blade leading edge and avoid acceleration and re-lamination of the turbulent boundary layer flow along the subsequent, suction-side curvature. Local disturbances in the boundary layer are produced which support the flow transition fixation.
  • the suction side of a conventionally designed leading edge with detached leading edge shock wave can be provided with a reflex to effect a rapid transition and the stabilization of the turbulent boundary layer flow. It is, however, also imaginable to combine this curvature on the suction side with a leading edge structure which effects the attachment of the leading edge shock wave.
  • FIG. 1 is a sectional view of a compressor blade with modified leading edge in accordance with a first embodiment of the invention, including a graphical representation of the shock wave attaching to the leading edge of the compressor blade and of the compression shock acting on the suction side as well as the respective boundary layer flow,
  • FIG. 2 is a side view of a part of a compressor blade, with the design of the leading edge differing from that of the blade shown in FIG. 1 ,
  • FIG. 3 is a side view of a part of the compressor blade shown in FIG. 2 , however with modified design of the leading edge of the latter,
  • FIG. 4 is a side view of still another embodiment of a compressor blade according to the present invention.
  • FIG. 5 is a sectional view of the compressor blade according to FIG. 1 , however with discontinuous curvature and corresponding boundary layer flow on the suction side.
  • the compressor blades 1 illustrated in FIGS. 1 to 4 feature in an upper area, in relation to the blade tip 12 , a specific form (2 to 5) of the blade leading edge 6 by which, as shown in FIG. 1 , attachment of the leading edge shock wave 14 immediately at the blade leading edge 6 and, consequently, the boundary layer flow on the suction side 13 of the compressor blade 1 shown in FIG. 1 is achieved in this blade area.
  • the blade leading edge has an elliptic (or also a parabolic) cross-sectional profile 2 , with the ratio of the two semi-axes a/b being smaller than or equal to 1/4. This means that the curvature radius of the leading edge is small and the blade thickness correspondingly low.
  • the cross-section of the compressor blade 1 is continuously curved.
  • a serration 3 is provided in an upper area of the blade leading edge 6 .
  • a boundary layer flow is here achieved which corresponds to that shown in FIG. 1 .
  • FIG. 4 still shows another variant of a compressor blade 1 with the inventive design of the leading edge 6 which features, in the upper area, a recessed, curved section 5 to provide a leading edge sweepback.
  • the resultant differences in incidence and the changed ratio between the incidence Mach number and the relative Mach number here again effect attachment of the shock wave to the blade leading edge 6 in the respective leading edge area.
  • the blade features a continuous, suction-side curvature and small thickness.
  • the design of the blade leading edge 6 in accordance with the variants 1 to 4 in combination with the attachment of the shock wave to the blade leading edge 6 results in the initially laminar boundary layer flow 7 changing into a turbulent boundary layer flow 9 at a transition point 8 located shortly downstream of the blade leading edge 6 and the flow is not accelerated beyond a certain degree, as a result of which the turbulent boundary layer flow 9 is not re-laminated, i.e. remains turbulent.
  • Vibration of the transition point 8 is, therefore, low and its distance to the compression shock 10 , which acts upon the suction side 13 of the compressor blade 1 and which is normally in the range of 45 to 70 percent of the blade width away from the leading edge 6 , is large enough to avoid, or at least limit, a vibration-augmenting reaction on the suction-side compression shock 10 .
  • the compressor blade 1 with the respective, particular design of the blade leading edge 6 features a continuous curvature on the suction side 13 which extends from the leading edge. It is, however, also imaginable that the curvature on the suction side of the compressor blade 1 , as shown in FIG. 5 , is designed with a reflex, in which case a concave depression 11 is provided in the forward area of the compressor blade 1 . In the depression 11 , the pressure of the laminar boundary layer flow 7 along the suction side 13 is increased, as a result of which it transits into a turbulent boundary layer flow 9 at the transition point 8 so given.
  • the leading edge 6 can be designed conventionally, i.e. thicker or, respectively, with an ellipse ratio of a/b equal to 1/2, for example.
  • the discontinuously curved shape of the suction side can be used in combination with one of the variants illustrated in FIGS. 1 to 4 in order to even better stabilize the turbulent boundary layer flow and avoid re-lamination, thus preventing the transition point from moving periodically.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The compressor blades of an aircraft engine are, in at least one natural-vibration critical area, designed such that at the blade leading edge (6), the leading edge shock wave (14) attaches to the leading edge (6), as a result of which a laminar boundary layer flow (7) on the suction side (13) quickly transitions into a turbulent boundary layer flow (9) which is kept constant and prevented from re-lamination by the further, continuous curvature of the suction side. Therefore, the transition, whose periodic movement is also suppressed, cannot communicate with a suction-side compression shock (10), preventing the compression shock from augmenting the natural vibrations of the blade occurring under certain flight conditions. The blade leading edge can, for example, be designed as an ellipse with a semi-axis ratio equal to or smaller than 1:4.

Description

This application claims priority to European Patent Application EP 04090215.7 filed Jun. 2, 2004, the entirety of which is incorporated by reference herein.
BACKGROUND OF THE INVENTION
This invention relates to a compressor blade with reduced aerodynamic blade excitation, in particular for the fan of aircraft engines.
Compressor blades with a large chord length used on certain types of aircraft-engine compressors are loaded to such an extent under certain operating conditions that damage to or life reduction of the compressor blade and the compressor disk may occur.
In order to reduce the stressing of the compressor blades, the assignee of the present application has previously filed a U.S. patent application on Apr. 5, 2004, to Erik Johann, entitled “COMPRESSOR FOR AN AIRCRAFT ENGINE”, in which it was disclosed to provide a flow transition fixation means in the form of a surface roughness applied to the suction side of the blade at a certain distance from and parallel to the blade leading edge and clearly ahead of the area of the compression shocks acting upon the blade. The local fixation of the transition area away from the compression shock area suppresses, or at least limits, coupling of the interrelated oscillation in the transition and compression shock area and, ultimately, avoids oscillation of the compression shocks or limits them to such an extent that their vibration-augmenting reaction on the natural vibrations of the compressor blade occurring under certain flight conditions, which eventually may lead to damage of the compressor blades and the compressor disk, is prevented.
BRIEF SUMMARY OF THE INVENTION
A broad aspect of this invention is to provide a design of compressor blades such that, independently of the already suggested fixation of the flow transition point by applying a surface roughness to the suction side, oscillation of the compression shocks (suction side) and their vibration-augmenting coupling to the natural vibrations of the blade is avoided or reduced, in order to avoid damage to the compressor disks and blades occurring under certain operating conditions.
It is a particular object of the present invention to provide solution to the above problems by a compressor blade designed in accordance with the features described herein. Further variants and—if combined—advantageous embodiments of the present invention will become apparent from the description below.
The essence of this invention is the particular design of the blade leading edge such that the leading edge shock wave effective there will not detach from, but rather attach to the leading edge to effect the transition from the laminar to the turbulent boundary layer flow at a short distance from the leading edge without accelerating it along the continuous curvature of the suction side starting at the leading edge, thus keeping the boundary layer flow constant without re-lamination. Accordingly, the transition point, whose periodic movement is largely suppressed and which is sufficiently remote of the compression shock on the suction side, can neither communicate with nor have a vibration-augmenting effect on the latter. Thus, reaction of the vibrations of the compression shock onto the critical natural vibrations occurring under certain flight conditions is suppressed or at least limited to such an extent that the natural vibrations are not augmented, or are not augmented to such an extent that overstressing in excess of the strength limits, with consequential material damage to the compressor blades and/or the compressor disk, occurs.
The shape of the blade leading edge can, for example, be a very slender ellipse or parabola—with an ellipse ratio of the semi-minor axis to the semi-major axis smaller than or equal to 1:4—with the curvature extending continuously on the suction side. As a result of this design, the leading edge shock wave attaches to the leading edge, the initially laminar boundary flow quickly transforms, i.e. at a short distance from the blade leading edge, into a turbulent boundary flow without being accelerated and re-laminated, and the transition point is thus not moved periodically.
To achieve the above effects, the leading edge can, in the upper, natural-vibration critical area, also have a serrated structure, pocket-type depressions or a recessed area acting as sweep-back. Additionally, other designs of the blade leading edge are possible to ensure attachment of the leading edge shock wave to the blade leading edge and avoid acceleration and re-lamination of the turbulent boundary layer flow along the subsequent, suction-side curvature. Local disturbances in the boundary layer are produced which support the flow transition fixation.
In accordance with still another feature of the present invention, the suction side of a conventionally designed leading edge with detached leading edge shock wave can be provided with a reflex to effect a rapid transition and the stabilization of the turbulent boundary layer flow. It is, however, also imaginable to combine this curvature on the suction side with a leading edge structure which effects the attachment of the leading edge shock wave.
BRIEF DESCRIPTION OF THE DRAWINGS
The present invention is more fully described in light of the accompanying drawings showing a preferred embodiment. In the drawings:
FIG. 1 is a sectional view of a compressor blade with modified leading edge in accordance with a first embodiment of the invention, including a graphical representation of the shock wave attaching to the leading edge of the compressor blade and of the compression shock acting on the suction side as well as the respective boundary layer flow,
FIG. 2 is a side view of a part of a compressor blade, with the design of the leading edge differing from that of the blade shown in FIG. 1,
FIG. 3 is a side view of a part of the compressor blade shown in FIG. 2, however with modified design of the leading edge of the latter,
FIG. 4 is a side view of still another embodiment of a compressor blade according to the present invention, and
FIG. 5 is a sectional view of the compressor blade according to FIG. 1, however with discontinuous curvature and corresponding boundary layer flow on the suction side.
 DETAILED DESCRIPTION OF THE INVENTION
The compressor blades 1 illustrated in FIGS. 1 to 4 feature in an upper area, in relation to the blade tip 12, a specific form (2 to 5) of the blade leading edge 6 by which, as shown in FIG. 1, attachment of the leading edge shock wave 14 immediately at the blade leading edge 6 and, consequently, the boundary layer flow on the suction side 13 of the compressor blade 1 shown in FIG. 1 is achieved in this blade area. This means that the situation shown in FIG. 1 applies similarly to the compressor blades 1 shown in FIGS. 2 to 4.
According to the blade variant shown in FIG. 1, the blade leading edge has an elliptic (or also a parabolic) cross-sectional profile 2, with the ratio of the two semi-axes a/b being smaller than or equal to 1/4. This means that the curvature radius of the leading edge is small and the blade thickness correspondingly low. The cross-section of the compressor blade 1 is continuously curved.
In the case of the blade variant according to FIG. 2, a serration 3 is provided in an upper area of the blade leading edge 6. Essentially, a boundary layer flow is here achieved which corresponds to that shown in FIG. 1.
According to the third variant shown in FIG. 3, several adjacent, pocket-style recesses 4 are provided at the blade leading edge 6 in lieu of the serration which essentially have the same effect as the above mentioned serration or the blade leading edge 6 with the specified ellipse ratio.
Finally, FIG. 4 still shows another variant of a compressor blade 1 with the inventive design of the leading edge 6 which features, in the upper area, a recessed, curved section 5 to provide a leading edge sweepback. The resultant differences in incidence and the changed ratio between the incidence Mach number and the relative Mach number here again effect attachment of the shock wave to the blade leading edge 6 in the respective leading edge area.
Also with the blade designs according to FIGS. 2 to 4, the blade features a continuous, suction-side curvature and small thickness.
The design of the blade leading edge 6 in accordance with the variants 1 to 4 in combination with the attachment of the shock wave to the blade leading edge 6 results in the initially laminar boundary layer flow 7 changing into a turbulent boundary layer flow 9 at a transition point 8 located shortly downstream of the blade leading edge 6 and the flow is not accelerated beyond a certain degree, as a result of which the turbulent boundary layer flow 9 is not re-laminated, i.e. remains turbulent. Vibration of the transition point 8 is, therefore, low and its distance to the compression shock 10, which acts upon the suction side 13 of the compressor blade 1 and which is normally in the range of 45 to 70 percent of the blade width away from the leading edge 6, is large enough to avoid, or at least limit, a vibration-augmenting reaction on the suction-side compression shock 10. Since the compression shock 10 cannot communicate with the boundary layer flow 7, 9 or the prevented periodic movement of the transition point 8 of the boundary layer flow, respectively, oscillation of the compression shock is limited to such an extent that the coupling effect between the natural vibrations (second and third flexural, first torsional) occurring at the compressor blade 1 under certain conditions (Reynolds number, Mach number, Strouhal number) and the compression shock vibrations is suppressed and, thus, the natural vibrations are prevented from exceeding a magnitude which causes damage to the compressor blades and the compressor disk.
According to the variants shown in FIGS. 1 to 4, the compressor blade 1 with the respective, particular design of the blade leading edge 6 features a continuous curvature on the suction side 13 which extends from the leading edge. It is, however, also imaginable that the curvature on the suction side of the compressor blade 1, as shown in FIG. 5, is designed with a reflex, in which case a concave depression 11 is provided in the forward area of the compressor blade 1. In the depression 11, the pressure of the laminar boundary layer flow 7 along the suction side 13 is increased, as a result of which it transits into a turbulent boundary layer flow 9 at the transition point 8 so given. Here, acceleration at the end of the depression 11 is so low that the turbulent boundary layer flow 9 remains constant and instabilities ahead of it are dampened. Basically, in the case of the discontinuously curved suction side of the compressor blade 1 as per FIG. 5, the leading edge 6 can be designed conventionally, i.e. thicker or, respectively, with an ellipse ratio of a/b equal to 1/2, for example. Preferably, however, the discontinuously curved shape of the suction side can be used in combination with one of the variants illustrated in FIGS. 1 to 4 in order to even better stabilize the turbulent boundary layer flow and avoid re-lamination, thus preventing the transition point from moving periodically.
LIST OF REFERENCE NUMERALS
    • 1 compressor blade
    • 2 blade leading edge structure: elliptic
    • 3 blade leading edge structure: serration
    • 4 blade leading edge structure: pocket-style recesses
    • 5 blade leading edge structure: recessed section
    • 6 blade leading edge
    • 7 laminar boundary layer flow
    • 8 transition point
    • 9 turbulent bounding layer flow
    • 10 compression shock, suction-side shock
    • 11 suction-side depression, reflex
    • 12 blade tip
    • 13 suction side
    • 14 leading edge shock wave

Claims (11)

1. A compressor blade for a gas turbine, comprising a suction and a pressure side and a blade leading and trailing edge the blade having a relatively long chord length, wherein the blade leading edge, at a blade tip area, operates at transonic and supersonic velocities, and, in at least one natural-vibration critical area, is structured such that a leading edge shock wave attaches to the blade leading edge, as a result of which an initially laminar boundary layer flow changes, at a transition point a short distance from the blade leading edge into a turbulent boundary layer flow which neither accelerates nor re-laminates within a continuous curvature extending from the blade leading edge on the suction side, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a compression shock on the suction side;
wherein the blade leading edge is structured as one of an elliptic or parabolic cross-sectional profile having an ellipse ratio (a:b) being equal to or smaller than 1 to 4 to effect attachment of the leading edge shock wave immediately to the blade leading edge.
2. A compressor blade in accordance with claim 1, wherein the blade leading edge includes an area having a serration to effect attachment of the leading edge shock wave.
3. A compressor blade in accordance with claim 1, wherein the blade leading edge includes an area having pocket-style recesses to effect attachment of the leading edge shock wave.
4. A compressor blade in accordance with claim 1, wherein the blade leading edge includes a recessed area acting as a sweep-back to effect attachment of the leading edge shock wave.
5. A compressor blade in accordance with claim 4, wherein the recessed area includes a concave curvature.
6. A compressor blade in accordance with claim 1, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant.
7. A compressor blade for a gas turbine compressor, comprising a leading edge, which at a blade tip area, operates at transonic and supersonic velocities, and a suction side which includes a concave depression near the leading edge to cause transition of a boundary layer flow from laminar to turbulent, and keep acceleration of the boundary layer flow low and flow of the turbulent boundary layer constant; the depression being positioned on the suction side between the leading edge and a midpoint of the blade, the depression being shallow with gradually tapering leading and trailing edges and extending across a substantial portion of the distance from the leading edge to the blade midpoint, the depression positioned to begin between the leading edge and a midpoint between the leading edge and the blade midpoint and having a width several times its depth, the blade having a relatively long chord length.
8. A compressor blade for a gas turbine, comprising a leading edge and a suction side, the blade having a relatively long chord length, the blade leading edge, at a blade tip area, operating at transonic and supersonic velocities, and the blade leading edge including an area having pocket-style recesses to attach a leading edge shock wave to the leading edge, and to transition a suction side boundary layer flow from laminar to turbulent at a point a short distance from the leading edge, with the turbulent boundary layer flow neither accelerating nor re-laminating within a continuous curvature extending from the leading edge, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a suction side compression shock, the pocket style recesses positioned so that a major portion thereof is positioned on the suction side of the blade adjacent the leading edge.
9. A compressor blade in accordance with claim 8, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant.
10. A compressor blade for a gas turbine, comprising a suction and a pressure side and a blade leading and trailing edge, the blade having a relatively long chord length, wherein the blade leading edge, at a blade tip area, operates at transonic and supersonic velocities, and, in at least one natural-vibration critical area, is structured such that a leading edge shock wave attaches to the blade leading edge, as a result of which an initially laminar boundary layer flow changes, at a transition point a short distance from the blade leading edge into a turbulent boundary layer flow which neither accelerates nor re-laminates within a continuous curvature extending from the blade leading edge on the suction side, thus suppressing a periodic movement of the transition point and preventing the transition point from communicating with a compression shock on the suction side;
wherein the blade leading edge includes a concavely curved recessed area acting as a sweep-back to effect attachment of the leading edge shock wave, the recessed area starting at a position radially inward from the blade tip and ending before a blade midpoint.
11. A compressor blade in accordance with claim 10, wherein the suction side is designed as a reflex which includes a depression in an area of the laminar boundary layer flow on the suction side in order to effect, by pressure increase, a transition into the turbulent boundary layer flow and keep acceleration of the boundary layer flow low and the boundary layer flow itself constant.
US10/870,437 2004-06-02 2004-06-18 Compressor blade with reduced aerodynamic blade excitation Expired - Fee Related US7484937B2 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EPEP04090215.7 2004-06-02
EP04090215A EP1607573B2 (en) 2004-06-02 2004-06-02 Compressor blade with reduced aerodynamic vibration induction

Publications (2)

Publication Number Publication Date
US20050271513A1 US20050271513A1 (en) 2005-12-08
US7484937B2 true US7484937B2 (en) 2009-02-03

Family

ID=34928806

Family Applications (1)

Application Number Title Priority Date Filing Date
US10/870,437 Expired - Fee Related US7484937B2 (en) 2004-06-02 2004-06-18 Compressor blade with reduced aerodynamic blade excitation

Country Status (4)

Country Link
US (1) US7484937B2 (en)
EP (2) EP2050929B1 (en)
CA (1) CA2493563C (en)
DE (2) DE502004010281D1 (en)

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100266428A1 (en) * 2008-01-07 2010-10-21 Suguru Nakagawa Propeller fan
US20100322779A1 (en) * 2007-07-11 2010-12-23 Suguru Nakagawa Propeller fan
US20130164488A1 (en) * 2011-12-22 2013-06-27 General Electric Company Airfoils for wake desensitization and method for fabricating same
EP2669475A1 (en) * 2012-06-01 2013-12-04 Techspace Aero S.A. S-shaped profile blade of axial turbomachine compressor, corresponding compressor and turbomachine
US20140356156A1 (en) * 2013-05-28 2014-12-04 Honda Motor Co., Ltd. Airfoil geometry of blade for axial compressor
WO2015138672A1 (en) * 2014-03-13 2015-09-17 Magna Electronics Inc. Vehicle cooling fan with aerodynamic stator struts
US20160003060A1 (en) * 2013-03-07 2016-01-07 United Technologies Corporation Hybrid fan blades for jet engines
US9488055B2 (en) 2012-06-08 2016-11-08 General Electric Company Turbine engine and aerodynamic element of turbine engine
US20170022820A1 (en) * 2015-07-20 2017-01-26 Rolls-Royce Plc Aerofoil
US20170198793A1 (en) * 2016-01-07 2017-07-13 Caterpillar Inc. Torque converters and methods for assembling the same
US20170261000A1 (en) * 2014-09-18 2017-09-14 Denso Corporation Blower
US20190078450A1 (en) * 2017-09-08 2019-03-14 United Technologies Corporation Inlet guide vane having a varied trailing edge geometry
US10253676B2 (en) 2013-12-20 2019-04-09 Magna Powertrain Bad Homburg GmbH Molded rotor for cooling fan motor
US10731469B2 (en) 2016-05-16 2020-08-04 Raytheon Technologies Corporation Method and apparatus to enhance laminar flow for gas turbine engine components
US12037115B2 (en) * 2011-08-19 2024-07-16 Aerovironment, Inc. Aircraft system for reduced observer visibility

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070231141A1 (en) * 2006-03-31 2007-10-04 Honeywell International, Inc. Radial turbine wheel with locally curved trailing edge tip
GB0620769D0 (en) * 2006-10-19 2006-11-29 Rolls Royce Plc A fan blade
GB0701866D0 (en) * 2007-01-31 2007-03-14 Rolls Royce Plc Tone noise reduction in turbomachines
US20100322774A1 (en) * 2009-06-17 2010-12-23 Morrison Jay A Airfoil Having an Improved Trailing Edge
DE102010023017A1 (en) 2010-06-08 2011-12-08 Georg Emanuel Koppenwallner Humpback whale blower, method for locally improving the flow in turbomachinery and vehicles
KR101346085B1 (en) 2011-12-23 2013-12-31 한국항공우주연구원 Transonic Compressor Rotor Design Method, And Transonic Compressor Rotor
US20130330184A1 (en) * 2012-06-08 2013-12-12 General Electric Company Aerodynamic element of turbine engine
US20140044556A1 (en) * 2012-08-07 2014-02-13 General Electric Company Last stage blade including a plurality of leading edge indentations
US9963974B2 (en) * 2012-10-23 2018-05-08 United Technologies Corporation Reduction of equally spaced turbine nozzle vane excitation
CN104612758A (en) * 2014-12-19 2015-05-13 中国民航大学 Low-pressure turbine blade with low loss
JP6504655B2 (en) * 2015-03-31 2019-04-24 三菱重工業株式会社 Ship propeller
CN104847697B (en) * 2015-05-18 2017-04-12 哈尔滨工程大学 Gas compressor with front edge blade of spheroid convex structure
CN105332948B (en) * 2015-10-23 2017-08-15 上海交通大学 A kind of implementation method of the bionical movable vane of compressor
CN105298924B (en) * 2015-10-23 2017-09-15 上海交通大学 Compressor bionics stator blade and its implementation based on humpback flipper
CN105626159A (en) * 2016-03-03 2016-06-01 哈尔滨工程大学 Variable geometry turbine with wavy concaved structures on front edges of movable blades
US10718214B2 (en) * 2017-03-09 2020-07-21 Honeywell International Inc. High-pressure compressor rotor with leading edge having indent segment
WO2018216255A1 (en) 2017-05-24 2018-11-29 株式会社Ihi Blade for fan and compressor
FR3073016B1 (en) * 2017-10-30 2019-10-18 Safran Aircraft Engines MODULATION OF THE SERRATIONS IN THE END OF DAWN
GB201817175D0 (en) * 2018-10-22 2018-12-05 Rolls Royce Plc Fluidfoil
US11441811B2 (en) * 2019-03-11 2022-09-13 Air Distribution Technologies Ip, Llc Undulated surface enhancement of diffuser blade for plenum slot diffuser
US11448418B2 (en) * 2019-03-11 2022-09-20 Air Distribution Technologies Ip, Llc Undulated surface enhancement of diffuser blades for linear bar diffuser
US11435111B2 (en) * 2019-03-11 2022-09-06 Air Distribution Technologies Ip, Llc Undulated surface enhancement of diffuser blades for round and rectangular ceiling diffuser
CN110821851B (en) * 2019-11-22 2021-03-09 南京航空航天大学 Multistage axial compressor expands steady structure based on sawtooth trailing edge blade
CN111622808B (en) * 2020-05-25 2021-05-04 武汉大学 Bionic blade based on steam turbine blade profile transformation and design method

Citations (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB347138A (en) 1930-01-22 1931-04-22 Charles Algernon Parsons Improvements relating to steam turbines
US1903823A (en) 1928-12-28 1933-04-18 Lougheed Victor Aerodynamic surface
GB750305A (en) 1953-02-05 1956-06-13 Rolls Royce Improvements in axial-flow compressor, turbine and like blades
US3304056A (en) 1965-03-19 1967-02-14 Hitachi Ltd Turbine blades
US3365126A (en) * 1965-09-01 1968-01-23 Gen Electric Compressor blade
US3403893A (en) * 1967-12-05 1968-10-01 Gen Electric Axial flow compressor blades
US3578264A (en) 1968-07-09 1971-05-11 Battelle Development Corp Boundary layer control of flow separation and heat exchange
GB2032048A (en) 1978-07-15 1980-04-30 English Electric Co Ltd Boundary layer control device
US5069043A (en) 1989-07-07 1991-12-03 Advanced Cooling Technology, Inc. Refrigeration system with evaporative subcooling
US5169290A (en) 1991-11-07 1992-12-08 Carrier Corporation Blade for centrifugal flow fan
US5209644A (en) * 1991-01-11 1993-05-11 United Technologies Corporation Flow directing element for the turbine of a rotary machine and method of operation therefor
US5540406A (en) 1993-10-25 1996-07-30 Occhipinti; Anthony C. Hydrofoils and airfoils
WO1998022711A1 (en) 1996-11-18 1998-05-28 Lm Glasfiber A/S The use of a turbulator for damping stall vibrations in the blades of a wind turbine
EP0937862A2 (en) 1998-02-20 1999-08-25 BMW Rolls-Royce GmbH Blade form for axial turbines
US6139258A (en) * 1987-03-30 2000-10-31 United Technologies Corporation Airfoils with leading edge pockets for reduced heat transfer
US6183197B1 (en) 1999-02-22 2001-02-06 General Electric Company Airfoil with reduced heat load
EP1081332A1 (en) 1999-08-31 2001-03-07 ROLLS-ROYCE plc Axial flow turbines
US6264429B1 (en) 1997-06-24 2001-07-24 Siemens Aktiengesellschaft Compressor blade or vane and compressor using a blade or vane
US6358012B1 (en) * 2000-05-01 2002-03-19 United Technologies Corporation High efficiency turbomachinery blade
US6358003B2 (en) * 1998-03-23 2002-03-19 Rolls-Royce Deutschland Ltd & Co. Kg Rotor blade an axial-flow engine
DE10054244A1 (en) 2000-11-02 2002-06-13 Honda Motor Co Ltd Turbine blade profile for axial turbine has inner curved surface and outer curved surface with leading and trailing edges and with specially positioned turning point to reduce shock wave
US6538887B2 (en) 2001-07-26 2003-03-25 Hewlett-Packard Company Fan blade providing enhanced performance in air movement

Patent Citations (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1903823A (en) 1928-12-28 1933-04-18 Lougheed Victor Aerodynamic surface
GB347138A (en) 1930-01-22 1931-04-22 Charles Algernon Parsons Improvements relating to steam turbines
GB750305A (en) 1953-02-05 1956-06-13 Rolls Royce Improvements in axial-flow compressor, turbine and like blades
US3304056A (en) 1965-03-19 1967-02-14 Hitachi Ltd Turbine blades
US3365126A (en) * 1965-09-01 1968-01-23 Gen Electric Compressor blade
US3403893A (en) * 1967-12-05 1968-10-01 Gen Electric Axial flow compressor blades
US3578264A (en) 1968-07-09 1971-05-11 Battelle Development Corp Boundary layer control of flow separation and heat exchange
US3578264B1 (en) 1968-07-09 1991-11-19 Univ Michigan
GB2032048A (en) 1978-07-15 1980-04-30 English Electric Co Ltd Boundary layer control device
US6139258A (en) * 1987-03-30 2000-10-31 United Technologies Corporation Airfoils with leading edge pockets for reduced heat transfer
US5069043A (en) 1989-07-07 1991-12-03 Advanced Cooling Technology, Inc. Refrigeration system with evaporative subcooling
US5313700A (en) 1991-01-11 1994-05-24 United Technologies Corporation Forming a flow directing element for a turbine
US5209644A (en) * 1991-01-11 1993-05-11 United Technologies Corporation Flow directing element for the turbine of a rotary machine and method of operation therefor
US5169290A (en) 1991-11-07 1992-12-08 Carrier Corporation Blade for centrifugal flow fan
US5540406A (en) 1993-10-25 1996-07-30 Occhipinti; Anthony C. Hydrofoils and airfoils
WO1998022711A1 (en) 1996-11-18 1998-05-28 Lm Glasfiber A/S The use of a turbulator for damping stall vibrations in the blades of a wind turbine
EP0954701A1 (en) 1996-11-18 1999-11-10 Lm Glasfiber A/S The use of a turbulator for damping stall vibrations in the blades of a wind turbine
US6264429B1 (en) 1997-06-24 2001-07-24 Siemens Aktiengesellschaft Compressor blade or vane and compressor using a blade or vane
EP0937862A2 (en) 1998-02-20 1999-08-25 BMW Rolls-Royce GmbH Blade form for axial turbines
US6358003B2 (en) * 1998-03-23 2002-03-19 Rolls-Royce Deutschland Ltd & Co. Kg Rotor blade an axial-flow engine
US6183197B1 (en) 1999-02-22 2001-02-06 General Electric Company Airfoil with reduced heat load
EP1081332A1 (en) 1999-08-31 2001-03-07 ROLLS-ROYCE plc Axial flow turbines
US6416289B1 (en) 1999-08-31 2002-07-09 Rolls-Royce Plc Axial flow turbines
US6358012B1 (en) * 2000-05-01 2002-03-19 United Technologies Corporation High efficiency turbomachinery blade
DE10054244A1 (en) 2000-11-02 2002-06-13 Honda Motor Co Ltd Turbine blade profile for axial turbine has inner curved surface and outer curved surface with leading and trailing edges and with specially positioned turning point to reduce shock wave
US6638021B2 (en) 2000-11-02 2003-10-28 Honda Giken Kogyo Kabushiki Kaisha Turbine blade airfoil, turbine blade and turbine blade cascade for axial-flow turbine
US6538887B2 (en) 2001-07-26 2003-03-25 Hewlett-Packard Company Fan blade providing enhanced performance in air movement

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
European Search Report dated Aug. 25, 2004.
European Search Report dated Sep. 17, 2004.

Cited By (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100322779A1 (en) * 2007-07-11 2010-12-23 Suguru Nakagawa Propeller fan
US8512004B2 (en) * 2007-07-11 2013-08-20 Daikin Industries, Ltd. Propeller fan
US8721280B2 (en) * 2008-01-07 2014-05-13 Daikin Industries, Ltd. Propeller fan
US20100266428A1 (en) * 2008-01-07 2010-10-21 Suguru Nakagawa Propeller fan
US12037115B2 (en) * 2011-08-19 2024-07-16 Aerovironment, Inc. Aircraft system for reduced observer visibility
CN104114815B (en) * 2011-12-22 2016-10-12 通用电气公司 Airfoil and corresponding manufacture method
US20130164488A1 (en) * 2011-12-22 2013-06-27 General Electric Company Airfoils for wake desensitization and method for fabricating same
CN104114815A (en) * 2011-12-22 2014-10-22 通用电气公司 Airfoils and corresponding fabricating method
US9249666B2 (en) * 2011-12-22 2016-02-02 General Electric Company Airfoils for wake desensitization and method for fabricating same
EP2669475A1 (en) * 2012-06-01 2013-12-04 Techspace Aero S.A. S-shaped profile blade of axial turbomachine compressor, corresponding compressor and turbomachine
US9957973B2 (en) 2012-06-01 2018-05-01 Safran Aero Boosters Sa Blade with an S-shaped profile for an axial turbomachine compressor
US9488055B2 (en) 2012-06-08 2016-11-08 General Electric Company Turbine engine and aerodynamic element of turbine engine
US20160003060A1 (en) * 2013-03-07 2016-01-07 United Technologies Corporation Hybrid fan blades for jet engines
US20140356156A1 (en) * 2013-05-28 2014-12-04 Honda Motor Co., Ltd. Airfoil geometry of blade for axial compressor
US9752589B2 (en) * 2013-05-28 2017-09-05 Honda Motor Co., Ltd. Airfoil geometry of blade for axial compressor
US10253676B2 (en) 2013-12-20 2019-04-09 Magna Powertrain Bad Homburg GmbH Molded rotor for cooling fan motor
US10337525B2 (en) 2014-03-13 2019-07-02 Magna Electronics Inc. Vehicle cooling fan with aerodynamic stator struts
WO2015138672A1 (en) * 2014-03-13 2015-09-17 Magna Electronics Inc. Vehicle cooling fan with aerodynamic stator struts
US20170261000A1 (en) * 2014-09-18 2017-09-14 Denso Corporation Blower
US10301942B2 (en) * 2015-07-20 2019-05-28 Rolls-Royce Plc Aerofoil
US20170022820A1 (en) * 2015-07-20 2017-01-26 Rolls-Royce Plc Aerofoil
US20170198793A1 (en) * 2016-01-07 2017-07-13 Caterpillar Inc. Torque converters and methods for assembling the same
US10731469B2 (en) 2016-05-16 2020-08-04 Raytheon Technologies Corporation Method and apparatus to enhance laminar flow for gas turbine engine components
US11466574B2 (en) 2016-05-16 2022-10-11 Raytheon Technologies Corporation Method and apparatus to enhance laminar flow for gas turbine engine components
US20190078450A1 (en) * 2017-09-08 2019-03-14 United Technologies Corporation Inlet guide vane having a varied trailing edge geometry

Also Published As

Publication number Publication date
EP2050929B1 (en) 2009-10-21
EP1607573A1 (en) 2005-12-21
CA2493563A1 (en) 2006-07-20
EP1607573B2 (en) 2012-10-03
DE502004009528D1 (en) 2009-07-09
US20050271513A1 (en) 2005-12-08
EP2050929A1 (en) 2009-04-22
CA2493563C (en) 2012-09-11
DE502004010281D1 (en) 2009-12-03
EP1607573B1 (en) 2009-05-27

Similar Documents

Publication Publication Date Title
US7484937B2 (en) Compressor blade with reduced aerodynamic blade excitation
US7997872B2 (en) Fan blade
EP2718182B1 (en) The split blended winglet
US8591195B2 (en) Turbine blade with pressure side stiffening rib
US6638021B2 (en) Turbine blade airfoil, turbine blade and turbine blade cascade for axial-flow turbine
US7878759B2 (en) Mitigation of unsteady peak fan blade and disc stresses in turbofan engines through the use of flow control devices to stabilize boundary layer characteristics
US20090136354A1 (en) Blade
US20100124500A1 (en) Turbomachine having an unducted fan provided with air guide means
US6776582B2 (en) Turbine blade and turbine
JP2010530333A (en) Small wings
JP5342637B2 (en) Airfoil for an axial compressor that enables low loss in the low Reynolds number region
WO2008094207A3 (en) High performance low noise helicopter blade aerodynamic design
JPH01159401A (en) Compressor blade with asymmetric nose small piece
JP2004324646A (en) Method and device for supporting tip of airfoil structurally
US5209643A (en) Tapered propeller blade design
RU2502640C2 (en) Structure with asymmetric ledge for compression shock structure changing
AU2003287542B2 (en) Laminar flow wing for transonic cruise
US7056089B2 (en) High-turning and high-transonic blade
US20130287579A1 (en) Blade of a turbomachine, having passive boundary layer control
US20230174231A1 (en) Airfoil With Supersonic Wave-Tripping Structure
JP4086374B6 (en) Aircraft blade
WO2019150637A1 (en) Aircraft
JP4086374B2 (en) Aircraft blade
US20160115798A1 (en) Rotor blade with reduced acoustic response
WO2005068782A1 (en) Laminor/turbulent transition zone stabilization on a turbine vane

Legal Events

Date Code Title Description
AS Assignment

Owner name: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:JOHANN, ERIK;REEL/FRAME:015901/0128

Effective date: 20040622

STCF Information on status: patent grant

Free format text: PATENTED CASE

FPAY Fee payment

Year of fee payment: 4

FPAY Fee payment

Year of fee payment: 8

FEPP Fee payment procedure

Free format text: MAINTENANCE FEE REMINDER MAILED (ORIGINAL EVENT CODE: REM.); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

LAPS Lapse for failure to pay maintenance fees

Free format text: PATENT EXPIRED FOR FAILURE TO PAY MAINTENANCE FEES (ORIGINAL EVENT CODE: EXP.); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

STCH Information on status: patent discontinuation

Free format text: PATENT EXPIRED DUE TO NONPAYMENT OF MAINTENANCE FEES UNDER 37 CFR 1.362

FP Lapsed due to failure to pay maintenance fee

Effective date: 20210203